This application is based on and claims the priority under 35 U.S.C. xc2xa7119 of German Patent Application 199 45 586.4, filed on Sep. 23, 1999, the entire disclosure of which is incorporated herein by reference.
The invention relates to a thermal protection system, especially providing heat shielding for space vehicles, including a flexible insulation mat arranged on the outer skin of the space vehicle.
During ascent and during reentry of a space vehicle into the atmosphere of any heavenly body, e.g. not only during reentry into the Earth""s atmosphere but also during reentry into the respective atmospheres of moons and planets such as Mars or Venus, the space vehicle is subjected to extreme aerodynamic, aerothermic, aeroelastic, mechanical and acoustic loads. For this reason, the ascent and reentry phases of a space mission are regarded as critical and determinative of the success of the mission. Therefore, it is important to equip the cold structure of the space vehicle with an effective thermal protection that can withstand all of the above described loads during the ascent and reentry phases, in order not to jeopardize the success of the entire space mission.
Various so-called reusable space transport vehicles are known, such as the U. S. space shuttle, the Russian BURAN, the originally planned European project HERMES, and the future Japanese space glider HOPE. For such reusable space transport vehicles, it is known to arrange ceramic tiles or shingles and other hot structures in the form of a rigid thermal protection system on the highly thermally loaded areas of the respective space vehicle, such as the underside or belly, the nose area, and the leading edges of the wings of the vehicle. On the other hand, the leeward sides of the components of the vehicle, which are not as highly thermally loaded, are typically provided with flexible insulation mats that are glued onto the cold structure of the space vehicle at such locations. Such conventional flexible insulation mats do not provide adequate thermal protection to be used at the highly thermally loaded areas mentioned above.
Throughout this application, the term xe2x80x9ccold structurexe2x80x9d refers to the structural components of the space vehicle, such as the skin thereof, which cannot directly withstand the high temperatures resulting during ascent and reentry. On the other hand, the term xe2x80x9chot structurexe2x80x9d refers to a structural component that can directly withstand the high temperatures arising during ascent and reentry.
In view of the above it is an object of the invention to provide a thermal protection system of the above described general type for a space vehicle, which achieves the advantages of a flexible insulation mat or the like, such as a low weight, a high flexibility for accommodating the different thermal expansions of the cold structure of the space vehicle and the hot outer wall or surface of the insulation system, and a simple and economical mounting and maintenance of the system on the space vehicle, not only for the leeward or relatively colder surfaces of the space vehicle, but also for the windward or relatively hotter surfaces of the space vehicle. It is a further object of the invention to reduce the overall cost of the total thermal protection system for the space vehicle, while providing a high degree of flexibility to accommodate higher degrees of deformation of the cold structure of the vehicle. Yet another object of the invention is to provide relatively thinner wall thicknesses of the insulation material and thereby achieve a considerable weight reduction of the overall thermal protection system. The invention further aims to avoid or overcome the disadvantages of the prior art, and to achieve additional advantages, as are apparent from the present specification.
The above objects have been achieved in a thermal protection system according to the invention, comprising at least one flexible insulation mat that is applied over a large surface area of the outer skin of the space vehicle. Particularly according to the invention, the thermal protection system further comprises a cover layer including at least one layer or ply comprising a ceramic fiber composite material, and a coating comprising an inorganic material applied onto the cover layer.
The particular composition of the fiber composite material of the cover layer can be adapted to the respective requirements at hand, which depend on the respective particular mission as well as the particular location at which the thermal protection system is being applied to the space vehicle. For example, the fiber composite material preferably and advantageously consists of inorganic fibers, which are preferably oxides, carbides or mixtures thereof, and which are in turn embedded in a matrix of the same material composition as the fibers. Depending on the particular selected composition, different particular thermal protection results as well as mechanical characteristics will be achieved.
The present thermal protection system according to the invention is especially further characterized in that it reliably prevents the penetration or permeation of hot gases into the insulation mat that is applied to the outer surface of the space vehicle, because especially the cover layer and coating form a pressure-tight or gas-tight outer skin covering the surface of the mat and thus of the space vehicle. Simultaneously, the cover layer and coating of the thermal protection system achieves a large surface distribution or spreading and introduction of the arising aerothermic, mechanical and acoustical loads, which are then conducted further through the underlying insulation mat and are then distributed and introduced over a large surface area of the cold structure of the space vehicle. Moreover, the protective layer provides an extremely effective protection against corrosive wear mechanisms and against the occurrence of hot gas oxidation. Also, by appropriately selecting the material of the coating that is provided on the thermal protection system according to the invention, the thermo-optical characteristics, and especially the absorption/emission characteristic of the surface of the thermal protection system, can be optimally adjusted. As a further advantage, the thermal protection system according to the invention is characterized by a rather low catalytic action in comparison to the prior art.
The application of the cover layer onto the insulation mat is preferably achieved according to the invention by means of adhesive bonding or gluing, sewing or stitching, or by a button or snap connection. The connection of the thermal protection system onto the cold structure of the space vehicle is advantageously achieved by an adhesive bonding, and especially by means of a high temperature resistant adhesive that cures at or under normal room temperature (e.g. under 30xc2x0 C.), and most preferably an adhesive based on silicone.